The present invention relates generally to gas turbine engines, and, more specifically, to casting of superalloy turbine blades therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in turbine stages for powering the compressor and performing external work, such as powering a fan in a turbofan aircraft gas turbine engine application.
Each turbine stage includes a stationary turbine nozzle having a row of nozzle vanes which discharge the combustion gases into a corresponding row of turbine rotor blades. Each blade includes an airfoil extending radially outwardly in span from an integral platform defining the radially inner flowpath boundary. The platform is integrally joined to a supporting dovetail having corresponding lobes mounted in a dovetail slot formed in the perimeter of a supporting rotor disk.
The first stage turbine blade first receives the hot combustion gases from the combustor through the corresponding first stage turbine nozzle. The turbine blades are typically hollow with internal cooling circuits therein specifically configured for cooling the different portions of the airfoil against the different heat loads from the combustion gases flowing thereover during operation.
The turbine airfoil includes a generally concave pressure side and circumferentially opposite, generally convex suction side which extend radially in span from a root at the platform to a radially outer tip, and extend axially in chord between opposite leading and trailing edges. The airfoil has the typical crescent radial profile or section which rapidly increases in thickness aft from the leading edge to the maximum width or hump region of the airfoil, which then gradually tapers and decreases in width to the relatively thin trailing edge of the airfoil.
The internal cooling circuits have numerous configurations, all of which share relatively long and slender radial cooling channels extending from root to tip, and separated chordally by typically imperforate and solid radial partitions bridging the opposite sidewalls of the airfoil.
The various internal cooling channels may be typically arranged in serpentine configurations over the midchord region of the airfoil. The leading and trailing edge regions of the airfoil typically include dedicated cooling channels, commonly having perforate partitions for effecting internal impingement cooling therein.
The various airfoil cooling channels have corresponding inlet channels extending upwardly through the blade platform and dovetail from the base thereof for receiving a portion of pressurized air bled from the compressor during operation. The channels have various outlets in the airfoil sidewalls, such as film cooling holes.
Since the turbine blades rotate with the supporting rotor disk during operation, they are subject to substantial centrifugal loads and corresponding stress. The centrifugal loads increase radially inwardly from the airfoil tip and are carried through the mounting dovetail into the supporting rotor disk.
Furthermore, each turbine blade is subject to the different heat loads over the radial span and axial chord thereof, and over the opposite pressure and suction sides, which correspondingly require different internal cooling for minimizing the amount of air bled from the compressor to correspondingly increase engine efficiency during operation.
In view of these exemplary mechanical and thermal operating conditions of the rotating turbine blades, they are typically manufactured by casting of superalloy metals having enhanced strength at the elevated temperatures of operation for ensuring a long useful life of the individual blades. Typical turbine blade materials include nickel or cobalt based superalloys which require corresponding casting to form the intricate interior and exterior shapes of the turbine blades, typically in state-of-the-art casting methods for achieving directional solidification of the casting grains or single crystal metallurgical configurations.
The common casting process is the lost wax method which begins with the fabrication of an intricate ceramic core that defines the various internal voids or flow channels of the turbine blade from the dovetail to the airfoil tip. The dovetail typically includes two or three relatively wide and short inlet channels, whereas the airfoil typically includes many more relatively narrow and long flow channels.
Accordingly, the ceramic core has a corresponding number of relatively long and slender ceramic legs that define the various long cooling channels in the airfoil, and a relatively few short and wide supporting stems that define the short inlet channels extending through the dovetail. The legs and stems are suitably joined together where appropriate, and are typically grouped together at the airfoil tip and grouped together at the base of the dovetail for providing an integrated and interconnected ceramic core assembly having suitable strength.
In the lost wax casting method, the core is initially positioned inside a pair of master dies which accurately define the three dimensional (3D) configuration of the turbine blade from dovetail to airfoil tip. Spaces or voids are provided between the core and the surrounding wax dies which represent the external configuration of the resulting metal in the later cast metal blade.
However, the voids between the core and the dies are initially filled with wax that is solidified into the 3D blade shape. The dies are removed, and the so cast wax version of the blade is surrounded in a ceramic slurry shell which is suitably cured hard.
The wax is melted and removed from the hardened shell again leaving the voids between the ceramic core and the so formed shell.
The voids are then filled with molten metal which is suitably solidified either directionally or in single crystal configuration as required for the superalloy metal being used to cast the individual blade.
The ceramic shell and ceramic core are then suitably removed, by chemical leaching for example, to leave behind the so cast superalloy metal blade. The solid cores are replaced by the hollow internal channels of the cooling circuits. The voids between the core legs and stems are replaced by the superalloy metal which structurally bridges together the opposite pressure and suction sides of the airfoil and corresponding opposite sides of the dovetail.
Since the resulting cast turbine blade has a 3D configuration varying significantly over its radial span, axial chord, and circumferential width, the internal cooling circuits have even more complex configurations in view of the multiple legs thereof separated by the corresponding radial partitions, with a relatively thin sidewall defining the external perimeter of the turbine blade.
The ceramic core which must be then specifically configured for the complex 3D turbine blade is yet even more complex in configuration in view of its relatively large length to width and thickness correspondingly creating long and slender ceramic legs. These slender legs are quite brittle in view of the high strength ceramic material used to withstand the high temperature of the molten metal during the casting process.
Accordingly, the ceramic core is quite fragile and is subject to inadvertent breakage during manufacture of the core itself, during handling of that core, and during the casting process. Core yield is a fundamental design parameter in the casting method.
Statistically, a small percentage of fabricated cores will break during the fabrication and use thereof and lead to a corresponding increase in cost of the manufacturing process. A core broken before the casting process is typically scrapped. A core broken during the casting process will typically lead to scrapping also of the so cast turbine blade.
Accordingly, ceramic core design includes features for strengthening thereof which are inherently limited by the final configuration of the turbine blade being cast. Thicker or stouter ceramic cores correspondingly result in thinner and weaker turbine blades in view of the loss in resulting cast metal. The external 3D configuration of a turbine blade is limited by the desired aerodynamic performance and efficiency of the engine, and simply making cores larger or blades larger decreases efficiency and performance of the turbine blade.
However, it is known to bridge together the core legs at the airfoil tip with a common block of ceramic. And, to bridge together the core stems at the base of the dovetail with another ceramic block. Furthermore, small ceramic ties may be preferentially located between the slender ceramic legs for bridging them together to increase the collective strength, but the ties necessarily result in a flow passage in the resulting turbine airfoil which may be undesirable for internal cooling performance.
It is also conventional to include individual ceramic ligaments between the relatively stout core stems and one or more of the slender core legs for typically supporting the legs of a serpentine cooling circuit which necessarily are joined together end-to-end near the airfoil tip and root in the final blade.
The individual ligaments are typically relatively long and slender with minimum sectional area to provide a temporary reinforcement in the ceramic core at the expense of the loss of supporting metal in the final turbine blade. After casting, the ligaments form a small flow channel between the dovetail inlet and the corresponding leg of the serpentine circuit, and a small metal ball is commonly brazed therein to block that undesirable flow communication in the final turbine blade.
The voids created by the core ligaments therefore interrupt the mechanical continuity and strength of the turbine blade typically in its shank region between the dovetail lobes and blade platform. This loss in supporting material in the final blade is a necessary compromise between reducing blade strength and increasing the strength of the fragile ceramic core to increase the yield thereof in the casting process for reducing manufacturing cost of the blades.
Yet further complicating the design of turbine blades is the ever increasing complexity of the internal cooling circuits, and the number of individual radial cooling channels therein, which correspondingly require even thinner and more slender ceramic legs in the ceramic core. In one advanced turbine blade undergoing study, the numerous slender ceramic legs are additionally strengthened by using slender core ligaments, but those slender ligaments are themselves subject to breakage, which also renders the ceramic core unusable.
Accordingly, it is desired to provide an improved ceramic core for a turbine rotor blade having slender internal cooling channels therein.